1. Field of the Art
The present invention relates generally to a vortex flow field and the apparatus and method to produce and sustain it and more particularly to a hybrid rocket engine and a method of propelling a rocket utilizing such vortex flow field. The flow field in accordance with the present invention is capable of providing separate regions or zones within and among one or more flowing fluids contained within a common chamber, without the need for diaphragms or other physical separators or barriers. It is evident and believed that the flow field of the present invention has utility to a wide range of applications. One general field of application is that of combustion chambers, and more particularly, that of combustion chambers and methods for rocket engines or the like and hybrid rocket propulsion systems. A combustion chamber and method in accordance with one embodiment of the present invention utilizes the unique flow field of the present invention to improve hybrid rocket fuel regression and increase mixing length in a rocket or other engine. Another embodiment is in the form of liquid rocket engine to prevent hot combustion products from contacting the chamber wall.
2. Description of the Prior Art
Virtually countless applications exist for a flow field which is compact and is capable of providing one or more separate regions or zones of flowing fluids within a container, without substantial mixing and without the need for any physical barrier or other separators between such regions or zones. With such a flow field, a chemical reaction, such as combustion, can be induced to incur in one region or zone while a separate fluid or process occupies another region or zone.
Many devices depend upon vortex flows for their successful operation, such as combustion chambers, cyclone separators, classifiers and the like that are in common use. All of these devices introduce swirling flow at one end of a passageway in which the flow follows a generally helical path to exit at the opposite end. Such conventional vortex flows do not achieve zonal separation as does the unique flow field that is the subject of the present invention.
Although the flow field in accordance with the present invention has significant applications in a variety of fields, it has particular application to the field of rocket engines and in one embodiment, specifically to hybrid rocket engines. Hybrid rocket engines denote a class of rocket propulsion systems in which one propellant is in fluid form and the other propellant is in the form of a solid grain. Typically, the fluid propellant is the oxidizer and the solid grain is the fuel. The oxidizer such as liquid oxygen is sprayed into the combustion ports in the solid fuel grain and caused to ignite. The hot combustion products sustain the combustion process until either the oxidizer flow is shut off or the fuel grain is depleted. In virtually all contemporary hybrids of today the limiting design factor is the rate at which the solid fuel grain can be caused to burn. The burn rate, often expressed as regression rate, is the rate at which fuel can be induced to vaporize or ablate off the grain surface so it can participate in the combustion process and contribute to rocket thrust. Because the rate is typically slow, conventional hybrid fuel grains must be made with large exposed surface areas. This is accomplished by casting large open combustion ports in the grain. The large ports waste volume in the high pressure casing, so that a larger, heavier, and more expensive case is needed than would be required if the fuel grain combustion ports could be much smaller by means of a flow field which improves the regression rate.
In recent years, hybrid rockets have received increasing attention from the National Aeronautics and Space Administration (NASA) sectors, Department of Defense, industrial aerospace participants and research institutions because their unique operational characteristics are capable of providing safer, lower-cost avenues to space than conventional solid propellant and liquid bi-propellant rocket propulsion systems. For example, hybrid rocket engines can be easily throttled for thrust tailoring, to perform in-flight motor shutdown and restart and to incorporate non-destructive mission abort modes. Also, since fuel in a hybrid rocket engine is stored in the form of a solid grain, such engines require only half the feed system hardware of liquid bi-propellant engines. Still further, the commonly used butadiene-based solid grain fuels are benign and neither toxic nor hazardous for storage and transportation. The hybrid solid fuel grain is also not susceptible to cracks and imperfections that can lead to catastrophic failure in solid rocket motor propellant grains.
However, despite these benefits, classical hybrid rocket engines, in which the oxidizer gas is injected into the combustion chamber at the end opposite the exit nozzle and in a direction parallel to the solid fuel grain, have not yet found widespread use for either commercial or military applications. Reasons for this include the fact that they suffer from relatively slow solid fuel regression rates, low volumetric loading and relatively poor combustion efficiency. For example, polymeric hybrid fuels such as hydroxyl-terminated polybutadiene (HTPD) regress generally about an order of magnitude slower than solid rocket motor propellants. In an effort to overcome these lower regression rates, complex cross-sectional geometries of the hybrid solid grain fuel with large wetted surface areas are often employed to achieve a large mass of flow rate of pyrolyzed vapor from the fuel grain. It has been shown that a three to fourfold increase in fuel regression rate can result in significant cost reductions, simplified grain manufacturing and large reductions in rocket inert weight.
In addition to problems associated with the low regression (fuel burning) rates of hybrid engines, the short straight line travel of the pyrolyzed fuel grain vapor and oxidizer as they traverse the combustion region results in incomplete mixing. This often necessitates the use of secondary combustion chambers at the end of the fuel grain to complete the combustion process. These secondary chambers add length and weight to the overall design and have the additional disadvantage of serving as a potential source and location of combustion instability.
Furthermore, both conventional hybrids and solid rocket motors must provide insulation layers between the solid propellant grain and the high pressure casing wall. This is necessary to prevent the exposure of the casing to the high temperature combustion gases when the grain material has been burned away out to the casing and no longer provides protection. The insulation adds weight and cost to the motor.
Accordingly, there is a need in the art for a flow field, and a structure and method for producing and sustaining it, which provides separate regions or zones of flowing fluids within a chamber. There is also a need in the art for a combustion chamber and method utilizing such a flow field, and particularly a combustion chamber and method for a hybrid rocket engine, which significantly increases the regression rate of the solid fuel grain and the effective chamber length and mixing within the combustion chamber. There is also a need for a combustion chamber and method utilizing such a flow field that prevents the hot combustion products from reaching the chamber wall.